Gas turbine engines are widely used and well developed generators of power used primarily for aircraft propulsion. An overview of gas turbine engine operation includes the compression of air to a high pressure high temperature condition in a compressor, injection and combustion of fuel with the compressed air and the expansion of the products of combustion through a turbine section which extracts a substantial amount of the energy present in the products of combustion. The turbine section powers the compressor section and in some instances also provides usable external mechanical power.
In large gas turbine engines the compressor section and turbine sections are of axial flow design and each stage comprises a disk having plurality of airfoils mounted on its rim. The blades and the disk move, the disk being mounted on a rotating shaft, and the disk blade assemblies are subject to very severe environments.
Historically blade and disk assemblies have been produced from separate components wherein a disk has the blades mechanically attached thereto. While this advantageously permits blade and disk to be of different materials, it adds substantially to the weight of the assembly relative to a unitary, integrally bladed rotor assembly.
The increases in the performance requirements for gas turbine engines, particularly military gas turbine engines, are leading to the introduction of integrally bladed rotors, wherein the blades are an integral part of the rotor and are either formed integrally with the disk or are metallurgically bonded to the disk. This reduces the weight deficit attributable to the prior art mechanical joining schemes.
A disk and blade assembly might typically comprise a single disk with about 100 blades attached thereto. In the prior art method of assembly employing mechanical joining techniques, it was relatively easy and straight forward to replace damaged blades simply by removing the damaged blade and replacing it with an identical undamaged blade. In the new environment of integrally bladed rotors, such repairs are far from simple. In the compressor and turbine sections of gas turbine engines, the blades and disks are operated at the outer limits of their property capabilities both in terms of stress and temperature. This means that any repair technique must produce repairs which have the strength of the parent metal, usually the blade material.
The need for repairs can arise both in service and in the initial fabrication of disk and blade assemblies. In service it is obvious that damage can arise from a variety of sources and also during the course of fabrication, the odds are fairly good that a statistically significant number of blades will have some defect.
There is also a need for a method to initially fabricate integrally bladed rotors by bonding individual blades onto a disk. Consequently, it is an object of the invention to repair or fabricate integrally bladed rotors.
Current blade bonding devices being developed address the bonding of the blades on a single integrally bladed rotor stage only. They use the philosophy of fixturing the rotor to a rigid frame, usually located on a mandrel which mates with the disk bore. The forces applied to the new blade being bonded are transmitted through the rig frame to the disk bore during the bond cycle. Although the forces are usually manageable (1,000 to 100,000 lbs.), an opportunity for elastic deflection of the frame and possibly the integrally bladed rotor itself exists, contributing to dimensional variances in the process.